The present invention relates generally to the design and construction of a speed reduction gearbox for transmission of power from a gas turbine engine to a helicopter rotor. More particularly, the present invention relates to a two-stage reduction main gearbox for a twin-engine helicopter.
To use gas turbine engines to drive helicopters, it requires speed reduction gearboxes between the engines and the helicopter rotors. This is true because gas turbine engines are high speed rotary equipment having components including an output shaft revolving at speeds from about 5,000 to 50,000 revolutions per minute. Many modern helicopters have two horizontally-spaced gas turbine engines with parallel, horizontally-spaced output shafts. To deliver the power from the output shafts of both engines to the main rotor, right-angle, speed-reducing "nose" gearboxes are used, each having an input shaft coupled to an engine output shaft. Each nose gearbox has an output shaft. A main gear reduction gearbox is coupled to the nose gearbox output shafts to harness the power from both engines, provide further speed reduction and increase output torque. Main gear reduction gearboxes include gear sets therein for reducing the shaft speed during the transmission of power from the nose gearboxes to the output devices. The transmission of power from the gas turbine engines to the output devices, which include the main rotor, the tail rotor, and various accessories, imparts substantial loads on the bearings and gears of the main reduction gearbox.
The application of gas turbine engines as a propulsion means for a helicopter often creates design parameter conflicts, such as the need for a durable long life gear train and the necessity to minimize the volume and weight of the transmission. Prior designers of gas turbine engine gear reduction gearboxes for helicopters have generally used multi-stage, main reduction gearboxes to effectuate significant shaft speed reduction. For smaller helicopters, two-stage reductions have been used. The first stage reduction has been limited to an input/output speed ratio of 5 to 1, i.e. 5:1. To the best of my knowledge, that has been the maximum first-stage reduction ratio considered acceptable in the helicopter industry. In efforts to deal with that limitation, and handle greater power inputs as developed by more powerful engines for larger helicopters, the typical practice has been to provide additional stages of gear reduction in the main gearbox. While a second stage has been acceptable for small helicopters, addition of a third stage has been the usual direction taken to handle larger helicopters with more powerful engines. Another approach has been to work within the generally-accepted 5:1 first stage reduction limit, by using a torque-splitting technique, but with attendant weight and cost penalties. Also, depending upon the proposed torque-splitting arrangement, failure to achieve true 50--50 torque splitting can result in overloading one of the two power paths, with attendant early gear failure.
The present invention is addressed to the continuing need for a simpler and lighter helicopter gearbox for handling more power.